Project Rigel: Mars Sample Return

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Project Rigel Mars Sample R
eturn


Kent Nebergall


1

Project Rigel: Mars Sample Return



MarsDrive Mars Sample Return Contest


Original:
November 1
, 2007

Illustrated: August 9, 2008


Kent Nebergall

Knebergall@
gmail.com




Abstract

This paper covers a design for Mars Sample Return entered in the MarsDrive Mar
s
Sample Return Competition.
The name Rigel was chosen because this star is the leading
footfall of the constellation Orion, and Orion is the key vehicle in the Vision for Space
Exploration.
Since this mission has a lot of nested criteria, it begins with

a survey of 20
possible configurations condensed into a series of graphs showing the limitations
imposed by these criteria.
After identifying these limits, the design with the greatest
margins is selected for two more design iterations. Schematics, comp
onent selections,
and mission timelines are then presented for this design.
Whenever possible, technology

and components
from other Mars missions
are

used. Where new equipment is designed,
it is relentlessly simplified to reduce the probability of mechan
ical failure and
development cost issues. Finally, budget and spin
-
off designs are considered t
hat would
build on this

design
. The spin
-
offs are considered very early in the design process to
allow for follow
-
up missions using as many components of the o
riginal mission as
possible, or to allow for changes in primary landing site based on future discoveries.


Project Rigel Mars Sample R
eturn


Kent Nebergall


2

Contents

Abstract

................................
................................
................................
...............................

1

Exami
ning the Solution Space

................................
................................
............................

3

Limitations of Current Entry, Descent, and Landing Technologies

...............................

3

The Solution Space Spreadsheet

................................
................................
.....................

4

Parameter Spreadsheet Assumptions

................................
................................
..........

5

Iteration 1 Solution Space Graphs and Conclusions

................................
.......................

9

Figure 1: Sample Size to Average Fuel Production Rate

................................
........

10

Figure 2: Sample Size to Solar Array Area

................................
.............................

11

Fig
ure 3: Sample Size to Tank Volumes

................................
................................
.

12

Figure 4: Sample Size to Vehicle Entry Mass (Scaled Rover)

................................

13

Figure 5: Sample Size to
Vehicle Entry Mass (Fixed 100 kg Rover)

......................

14

Figure 6: Sample Size to Vehicle Entry Mass (MER
-
sized 185 kg Rover)
.............

15

Design Figur
es for Iterations 1, 2, 3A, and 3B

................................
.............................

16

Iteration 3A and 3B Clarifications and Changes

................................
..........................

19

Examining Configuration Issues

................................
................................
.......................

21

General Design Parameters

................................
................................
...........................

21

The Vehicle Design
................................
................................
................................
...........

22

Sample Return Capsu
le

................................
................................
.............................

22

Earth Return Cruise Stage

................................
................................
.........................

28

Trans
-
Earth Injection (TEI) Stage (Second Stage)

................................
...................

32

Ascent (First) Stage
................................
................................
................................
...

37

Lander Components

................................
................................
................................
..

41

Launch Pad Components

................................
................................
..........................

46

Landing Site and Lander Operations

................................
................................
........

49

Sample Loading Bay and Gantry

................................
................................
..............

53

Lander Science Gear

................................
................................
................................
.

55

ISPP System

................................
................................
................................
..................

57

Daily ISPP Operation Cycles

................................
................................
....................

60

Cleaning the Array

................................
................................
................................
....

62

The Overall ISPP Production Cycle

................................
................................
.........

63

Solar Array

................................
................................
................................
................

65

Rover

................................
................................
................................
.........................

69

Mars
-
Bound Cruise Stage

................................
................................
.........................

73

Launch Vehicle

................................
................................
................................
.........

74

Biological Protection

................................
................................
................................

74

Avoiding a Single Point of Failure

................................
................................
...........

76

Testing
................................
................................
................................
.......................

77

Ground Operat
ions

................................
................................
................................
....

78

Mission Cost

................................
................................
................................
.................

80

Cost Factors Limiting Standard Models

................................
................................
...

80

Cost Estimation Tools [22]

................................
................................
.......................

81

MSR Follow
-
On Missions

................................
................................
................................

85

Mid
-
Latitude Variant

................................
................................
................................

85

Polar Ice Cap Variant

................................
................................
................................

86

Other Possible MSR
-
Derived Missions

................................
................................
....

88

Project Rigel Mars Sample R
eturn


Kent Nebergall


3

Gap Analysis and Iteration 4+

................................
................................
..................

89

Conclusion


Public Psychology and Mars Exploration
................................
...........

90

References

................................
................................
................................
.....................

91


Examining the
S
olution
S
pace


Limitation
s of
Current Entry, Descent, and Landing
Technologies


A recent NASA lecture
discussed the issue of Entry, Descent, and Landing (EDL) on
Mars. In a nutshell, the state of the art has depended on expensive testing from the
Viking era for parachute and entr
y capsule design to the point that with the Mars Science
Laboratory (MSL) mission, we will have reached the upper limit of our current
engineering for aeroshell design. Even validating existing designs requires launching
sounding rockets from high altitud
e balloons at a cost of $20 million per flight. Due to
ballistic coefficients, landing capsules cannot grow larger or heavier than MSL

for this
design
. Further, it was stated that while MSL plans to land an 800 kg payload on the
surface, a Mars Sample Re
turn (MSR) mission would need to land 1200 kg

[1]
. Note that
these numbers ignored the MSL landing stage, which would bring the total landed mass

for MSL
to 1319 kg, plus 219 kg propellant.



Parameter

Viking

[5]

MER

[7]

MSL

[9]

NASA MSR

Landing Ellipse

NA

80,000 m

20,000 m

20 m

[1]

Capsule
Diameter

3.5 m

2.65 m

4.5 m

4.5 m

[9]

Entry Mass

900 kg

800 kg

3250 kg

3250 kg

[9]

Landed Mass

576 kg

1
74

kg Rover

348 kg Lander

775 kg Rover

544 kg Descent Stage

219 kg Propellant

1200 kg

[1]



Viking
development
dovetailed the end of the Apollo
-
era budgets
, and was originally to
be launched on
a
Saturn
booster

under the name Voyager 1973 [16]
. There have been
statements from NASA scientists that Orion technology could be used for Mars Sample
Return

[2]
. While th
e Orion capsule is the same shape and is larger than the MSL
capsule (5 m as opposed to 4.5 m), the capsule is far too heavy as designed to land on the
surface of Mars. A ballistic coefficient of 100 is ideal for Mars entry, whereas
a
coefficient of
150 i
s too much.
The
Orion

CM ballistic coefficient

is 325

[1]
.
Theoretically, a lighter version could be created, but the commonality of parts would be
sacrificed.

Project Rigel Mars Sample R
eturn


Kent Nebergall


4

The Solution Space Spreadsheet

For the sake of completeness, this paper
initially
cover
s

twe
nty
strawman

design
s in a
spreadsheet

based on sample return size
. I
t

then selects a single design, iterates it with
more detailed information, and
expand
s

it into three variants optimized for operation at
different latitudes
.


Most of these designs will

exceed the criteria that the total mission cost

be held to $2

billion. That said, extending the solution space well beyond this

limit

for
the initial
solution space analysis
opens several doors.




While the scaling up of current technology to sample retur
n technology is a fairly
short jump, the further scaling towards the human
-
rated technology scale should
be examined to see i
f any dual
-
use technologies can be determined
, or if there is
a
suggestion

for a post
-
MSR mission that would help bridge the engine
ering gap
between MSR and human missions.





Many technologies, in particular methane engines, only currently exist at much
larger scales than required by a MSR mission. Given the high cost of engine
development
, including a larger scale mission within th
e data set is a logical
move

for the
initial analysis.




Certain technologies, such as life support
, can
use of Reverse Water Gas Shift and
related IS
PP methods
. By using these technologies in MSR, we increase
the
duration of testing under Mars conditions
before human missions

trust this
technology. Therefore including

these items, as well as large
scale power
production and other technologies, becomes a logical move for a forward
-
looking
program.


Showing all 3000 cells in the 20
-
design first iteration sp
readsheet is not physically
possible in this paper. Instead, we will first explore the assumptions from the first,
second, and third iterations in the table below. In the next section, several graphs will
show the data from the 20
-
design first iteration
to show how the core design was selected.
Finally, another table will list the selected first iteration design, the second
iteration,

and
the two variants of the third generation design.

Project Rigel Mars Sample R
eturn


Kent Nebergall


5

Parameter
Spreadsheet Assumptions

Below are the assumptions for
the three iterations of design discussed in this paper.


Parameter

Assumption

Data Source

Propellant

Ethylene and Oxygen

This combination is
ideal for smaller missions
due to the reduced need for hydrogen and
smaller tanks needed for the return vehicle.

Stage 1
Thrust/Mass
Ratio

0.7662

These ratios are from Mars Direct

[3]
. Th
e
first stage

also follows the general principle of
a rocket generating twice the trust required to
hover in local gravity.

A single engine is used
for both
a
scent and
d
escent, and

this method is
validated later in the paper.

Deeper work on this indicates the ratio of
0.894 is
ideal

for a liquid fueled first stage, but
that
0.7662 doesn’t induce much gravity loss
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Project Rigel Mars Sample R
eturn


Kent Nebergall


6

Stage 2 (TEI)
Dry Mass,


Stage 1
(Ascent) Dry
Mass

Engine mass: 15.5
percent of thrust.

Tank Mass:
12
percent of propellant
mass.

RCS mass: 7 kg per
pod including 2 kg
propellant.

Engine mass uses a 15.5:1 thrust to weight
ratio. This is
based on a study of existing
methane engines [17].

Tank mass uses a heavier version of the
standard formula becaus
e it is divided into two
sections per tank. Since the mass of the
propellant doesn’t change, the mass of the tank
doesn’t change regardless of shape, so the
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Project Rigel Mars Sample R
eturn


Kent Nebergall


7

circularization.

Non
-
rotating planet
us
ed so that design
will not be latitude
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dependant.

and drag of the remaining vehicle.

J C Whitehead
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fte牡瑩潮′
-
3BF

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-
牥摵湤d湴⁥瑨y汥le⽯/y来渠灲n摵d瑩潮⁵o
楴⁩渠
Dr. Zubrin’s original paper [4]


瑨t猠s汬潷猠o潲o
摵灬數⁲ 摵湤d湣y⁩渠桡 摷d牥.

Project Rigel Mars Sample R
eturn


Kent Nebergall


8

H2 Tank Dry
Mass

Assumes a standard
aircraft aluminum
alloy 5 mm thick.

This figure is for mass estimates alone, as the
real tank would have a metal inner shell
surroun
ded by wound composite for

both
strength and insulation, and finally insulation
layers.

ISRU
Electrical
Power Demand

The average kg of
propellant needed per
day (total propellant
divided by
432.8
days) multiplied by
575 watts per kg
needed.

While Dr. Zubr
in’s original work
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癡汵敳⁣牯獳r搠潶e爠a湤⁩猠牥污瑩癥ly⁡cc畲ute
景f⁰潷 爮

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day猠
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-
潮oy
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e
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p畲uace 䅲ea

㔳〠睡瑴
-
桯畲猠灥爠
p潬⁰敲㈮

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-
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灯楮p

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c潮獥牶r瑩癥 ca汣畬l瑩潮o.

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晲潭

fpo唠
摥浡湤⁴漠
p潬慲⁁ 牡y
獵灰sy.

m潷o爠摥浡湤‪‱㈠
桯畲猠⼠㔳〠睡瑴
-
桯畲猠灥爠卯r ㈮

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-
桯畲猠h㌶〰潵3e猩Ⱐ潲⁷a瑴
-
桯畲猠灥
爠獯氮†
䕸瑥湳楶t⁤楧g楮g⁩湴漠瑨敳o⁤ 獩g湳⁩浰m楥猠
瑨攠摥浡湤⁩猠扡獥搠潮d睡瑴
-
桯畲猠摵h⁴漠 桥
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潵瑰畴⁩猠扡獥搠潮⁳潬a爠潵瑰畴癥r⁡‱㈠桯畲
摡y⁛㑝Ⱐ獯⁩琠桡猠扥e渠浵n
瑩灬pe搠dy‱㈠瑯
浡瑣栠瑨攠潵瑰畴映瑨攠䵅M
-
瑥t桮潬潧y⁳潬a爠
a牲ay献

p潬慲⁁ 牡y
ta瑴
-
桯畲猠⼠ㄲ⸵‽†
This figure appears in one of Dr. Zubrin’s
Project Rigel Mars Sample R
eturn


Kent Nebergall


9

Mass

solar array kg

papers [4] and is verified by comparison with
other solar arrays for Mars applications [7
].

Landing
Aeroshell,
Heat Shield,
and Cruise
Stage Mass,
Shape, and
Volume

MER


㈮㘵2
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摥獩s渮†

䵓i


㐮㔠洠
摩慭d瑥t⁦潲 a汬
潴桥os

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潮oy⁨ 牤睡re⁡湤漠na浰me⁩猠ㄮ 㐵⁴4浥猠
瑨慴
映瑨f⁍䕒⁥湴 y慳献†䅳⁳畣栬⁴桩猠
灵獨p猠瑨攠扡汬楳i楣⁣潥晦ic楥湴i⁴桥⁣a灳畬e
潵琠潦⁴桥⁳ 潰o映睨w琠t桥⁳異 牳潮楣 条p
-
物湧⁰ 牡c桵瑥⁣a渠桡湤ne⁛ㅝ⸠.潲⁴桥⁳ 步映
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楳楳ie搠楮⁢潴栠䵅h⁡n搠䵓i⁥湴n
y ca灳畬e献†

o潶o爠䵡獳

㌠3潬畴楯渠ie瑳W

m牯灯r瑩潮o氠⠲l
-
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a湤⁦楸e搠d琠ㄸ㔠歧
⡍Eo
-
獩se搩

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瑨攠te獴映瑨攠癥桩c汥⸠⁔桥⁳浡汬e獴⁲潶s爠楳r
潶o爠瑷楣i⁴桥⁳楺e映卯 潵牮o爮

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yⰠ扥ca畳u映瑨f⁲ 瑩漠潦⁳o浰me
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-
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睩瑨⁤敭潮獴牡瑥搠摵ta扩汩ty⁩渠 桥⁍a牳r
e湶楲潮浥n琮

Iteration 1
Solution Space
Graphs and Conclusions

Iteration 1 assumed a perfect textbook vehicle with zero fuel waste. Issues with RCS
mass and using propellant for driving the pumps were not included in this iteration. That
said, some issues,
such as
duplex redundancy (and double

mass) for the ISRU systems
,

is
included in this estimate. This was the rough cut to determine the basic shape and size of
the vehicle.


W
ith

these charts
, there is

a flat
curve for the first three data points
. The first two points,
for 0 kg and 0.5 kg

return sample mass, respectively, assume that the vehicle is fit into
the Mars Exploration Rover (MER) 2.65 meter diameter aeroshell. The numbers after
that, starting with a repeat of the 0.5 sample, assume the use of the Mars Science
Laboratory (MSL) 4.
5 meter diameter aeroshell.

One of the first conclusions is that
the
MER capsules are not
practical for a MSR design, but that the MSL capsule has sufficient
room for fairly ambitious MSR designs.


Project Rigel Mars Sample R
eturn


Kent Nebergall


10

Figure 1: Sample Size to Average Fuel Production Rate

Propellant Production Per Sol (500 Days)
0.00
0.50
1.00
1.50
2.00
2.50
3.00
3.50
4.00
4.50
0.0
0.5
0.5
0.7
1.0
1.2
1.4
1.7
1.9
2.2
2.4
2.6
2.9
3.1
3.3
3.6
3.8
4.1
4.3
4.5
4.8
5.0
Return Sample Mass (kg)

These numbers are based on the assumption of a 500 day stay, with 432.8 days devoted to
producing ethylene and oxygen using the hydrogen on board, and the remaining days
devoted to producing
only
oxygen to reach the appropriate fuel

to oxygen

ratio.

The

durations of these phases were calculated each time and perfectly consistent regardless of
the scale of the vehicle. Note that a system producing 5 kg/day
would be big enough

for
life support on a human mission in terms of recycling carbon dioxide exhale
d by
crewmembers. A dual
-
use system with some level of redundancy (say two matched units
for Mars and three for a human
long
-
term vehicle
) could give field testing of these units
and economy of scale.


Project Rigel Mars Sample R
eturn


Kent Nebergall


11

Figure
2
: Sample Size to Solar Array Area

Solar Panel Area (Square Meters)
0
10
20
30
40
50
60
0.0
0.5
0.5
0.7
1.0
1.2
1.4
1.7
1.9
2.2
2.4
2.6
2.9
3.1
3.3
3.6
3.8
4.1
4.3
4.5
4.8
5.0
Return Sample Mass (kg)

The real

problem for scale is shown above. Even an empty sample capsule requires an
array of 18 square me
ters
.
Cost, ballistics, mass, and other technology issues
are less
limiting than the issue of how much solar array can deploy autonomously, in rough
terrain,

with a high likelihood of success, while

being

pack
ed

into the smallest, lightest
configuration possible.


Project Rigel Mars Sample R
eturn


Kent Nebergall


12

Figure
3
: Sample Size to Tank Volumes

Cumulative Tank Capacities (Cubic Meters)
0.00
0.50
1.00
1.50
2.00
2.50
3.00
3.50
0.0
0.5
0.5
0.7
1.0
1.2
1.4
1.7
1.9
2.2
2.4
2.6
2.9
3.1
3.3
3.6
3.8
4.1
4.3
4.5
4.8
5.0
Return Sample Mass (kg)
Stage 2 Fuel
Stage 2 Oxygen
Stage 1 Fuel
Stage 1 Oxygen
Hydrogen Tank

This graph gives the volume of each major tank in the system, starting with the hydrogen
tank and then the
first and second stages in turn. Values are cumulative but listed
individually


no combined tanks are
shown. If tanks were combined, or made into
a
single pressure vessel with a bulkhead in between, this could be made more efficient for
each stage. Als
o note these are interior
capacities, not e
xterior volumes
. This graph
proved that fitting a MSR vehicle in the 4.5 meter MSL capsule is realistic, even if the
capsule itself is not spherical and the tanks cannot be fused together in this manner.

Project Rigel Mars Sample R
eturn


Kent Nebergall


13

Figure

4
: Sample Size to Vehicle
Entry Mass

(Scaled Rover)

Entry Mass (kg) With Scaled Rover Compared to MER and MSL
0.0
500.0
1,000.0
1,500.0
2,000.0
2,500.0
3,000.0
0.0
0.5
0.5
0.7
1.0
1.2
1.4
1.7
1.9
2.2
2.4
2.6
2.9
3.1
3.3
3.6
3.8
4.1
4.3
4.5
4.8
5.0
Return Sample Mass (kg)
Nominal Entry Mass
MER/MSL Entry Mass
0.5 Sample, MER Capsule
15 kg Rover
0.5 Sample, MSL Capsule
15 kg Rover
3.3 kg Sample,
68.7 kg Rover
MSL Baseline
MER Baseline

The term “Scaled Rover” means that the rover mass is scaled to the mass of the overall
vehicle, with data points evenly spread from 20 kg on the low end to 100 kg on the high
end. For reference, Sojo
urner
is

10.6

kg
[5]
and the MER rovers are 1
74

kg

[7]
. It is
assumed that even without a rover, local sampling equipment
for drilling
would
still
be
roughly 20 kg.


This graph tells us several very important things. First, as noted, the first two figu
res use
the MER
2.65 meter capsule
, whereas all the rest use the MSL 4.5 m
eter

capsule. Note
how much heavier the vehicle is than the MER in terms of entry mass. The actual values
for the 0.5 kg sample are 1068.7 kg for the
strawman MSR

versus 820 kg for

MER.
Since MER was already nearly double the landed mass of the Pathfinder for the same
size
entry capsule, it is reasonable to assume the ballistic coefficient of the MER is already
close to maximum. To add another 45.7 percent to the mass is unrealist
ic. It would not
slow down enough with the heat shield to deploy the supersonic ring parachute.
Further,
even if it did work, the scale of the TMI mass is also increase
d

by 23.1 percent, which
would result in reengineering the launch vehicle interface to

include a larger TMI stage
and
l
arger launch vehicle.
Also note these are strawman figures for a mathematically
perfect design with no margins


the actual numbers will grow with future iterations.


When we jump to the much larger MSL 4.5 meter
entry veh
icle
, the relative entry mass
gives a great deal of payload flexibility and design margin.
All else being equal, the
Project Rigel Mars Sample R
eturn


Kent Nebergall


14

design can be expanded to return a 3.3 kg sample
return
before becoming
exceeding the
MSL
mass in the
ballistic coefficient. MSL is desig
ned to land at much higher altitudes
than earlier missions
. Also, MSR has the advantage of being able to handle larger
landing masses simply by adding more fuel


at least until the combined mass equals that
of MSL. This gives the MSL capsule


MSR missi
on design

combination

a very wide
range of capability. Given this range, the appropriate course is probably to build a
minimalist mission with a lot of safety margin for the first round, then build on the design
heritage for one or two follow
-
up missions.



A
ccording to the earlier graph, the solar array needed to fuel this 3.3 kg sample return
mission would be
40.78 square meters. Logically, a better use of this margin (at least at
this phase) would be to scale up the rover to increase the quality

and va
riety
of samples
rather than
increase the sample mass returned to scale up the

quantity.


Figure 5: Sample Size to Vehicle Entry Mass (Fixed 100 kg Rover)


For this illustration,
all the rovers are
simply scaled up to 100 kg ra
ther than making them
proportional to the remainder of the vehicle. Note that the
maximum
sample size now
drops from 3.3 kg (wi
th a 68.7 kg rover) to 3.1 kg. In other words, with the addition of
30.3 kg of rover mass, we only cut 0.2 kg from the sample r
eturn mass.




Entry Mass (kg) With 100 kg Rover

0.0

500.0

1,000.0

1,500.0

2,000.0

2,500.0

3,000.0

0.0

0.5

0.5

0.7

1.0

1.2

1.4

1.7

1.9

2.2

2.4

2.6

2.9

3.1

3.3

3.6

3.8

4.1

4.3

4.5

4.8

5.0

Retur
n Sample Mass (kg)

Nominal Entry Mass

MER/MSL Entry Mass

MSL Baseline

MER Baseline

0.5 Sample, MER Capsule

100 kg Rover

0.5 Sample, MSL Capsule

100 kg Rover

3.1 kg Sample,

100 kg Rover

Project Rigel Mars Sample R
eturn


Kent Nebergall


15

Figure 6: Sample Size to Vehicle Entry Mass (MER
-
sized
185 kg
Rover)


While the 100 kg rover above could be designed and may even use some parts from
MER, it seems logical to scale the mission as if the e
ntire rover were built from MER
comp
onents
.
Creating more
copies of the MER rovers with

sample return modifications
would involve
minim
al
developmental cost. Further, since the
vast majority of these
systems
have
been field
-
tested over two vehicles

for twice the required mission of the
MSR

rover, they can be assumed to be robust enough for this operation.

Note that the
NASA Fa
ct Sheet states that the rovers are

174 kg, whereas the web site states it weighs
185 kg. Since it will be modified for sample return anyway, the higher figure is us
ed.

Entry Mass (kg) With 185 kg (MER-Sized) Rover
0.0
500.0
1,000.0
1,500.0
2,000.0
2,500.0
3,000.0
0.0
0.5
0.5
0.7
1.0
1.2
1.4
1.7
1.9
2.2
2.4
2.6
2.9
3.1
3.3
3.6
3.8
4.1
4.3
4.5
4.8
5.0
Return Sample Mass (kg)
Nominal Entry Mass
MER/MSL Entry Mass
MSL Baseline
MER Baseline
0.5 Sample, MER Capsule
185 kg Rover
0.5 Sample, MSL Capsule
185 kg Rover
2.6 kg Sample,
185 kg Rover

Increasing the rover to MER proportions has dropped the maximum sample size to 2.6 kg


which would still require 35.16 square meters of solar array to power the In Situ
Propellant Production (ISPP) system. If the sample is scaled down to 0.5 kg, an
d the
rover remains the mass of the MER, we still have
a vehicle that is 73.1 percent the entry
mass and 77.5 percent the launch mass of MSL.
This is the initial, idealized iteration.
However, t
hese are very broad margins to work with.



The nominal mis
sion, returning a 0.5 kg sample to Earth, can carry an additional 739.7 kg
of equipment including the rover. This vehicle lands with first stage fuel tanks that are
20
-
30 percent filled to capacity (depending on how much equipment is lande
d and the
sample

return mass). As a working model, this uses the best of (by then) past missions
while developing only enough new hardware (sample collection equipment, return
vehicle,
ISPP,
large solar arrays) to complete the basic mission.

Project Rigel Mars Sample R
eturn


Kent Nebergall


16

Design Figures
for Iteratio
ns 1, 2,
3A
, and
3B

Iteration
1
is the selected design from the 20 listed above. Iteration 2 adds RCS to the
second stage and some additional margins. This is where the first release of this paper
ended. I
teration 3

adds more realistic waste margins

to
the system, propellant pump fuel
demands, and
some atmospheric factors
.
3A

assumes the first stage circularizes the orbit,
whereas
3B

assigns that task to the second stage.


MSR Project Rigel

Iter
.

1

Iter
.

2

Iter
.

3A

Iter
.

3B

Unit


Stage 1 Rocket Equat
ion






Propellant Mass

589.63

761.96

933.24

877.12

kg

Mass before burn (Mo)

873.94

1129.36

1357.84

1320.97

kg

Mass after burn (Mf)

284.30

367.40

424.60

443.85

kg

Gravity

9.81

9.81

9.805

9.805


deltaV (M/Sec)

4140

4140

4157

3900

M/sec

ISP

376.00

376
.00

364.7

364.7

sec


Stage 2 Rocket Equation






Propellant Mass

126.69

152.08

164.52

188.55

kg

Mass before burn (Mo)

196.33

235.68

250.60

277.15

kg

Mass after burn (Mf)

69.64

83.60

86.082

88.6

kg

Gravity

9.81

9.81

9.805

9.805


deltaV (M/Sec)

3821

3
821

3821

4078

M/sec

ISP

376.00

376.00

364.7

364.7

sec


Landing Rocket Equation






Propellant Mass

163.78

196.40

219.81

217.66

kg

Mass before burn (Mo)

1042.63

1250.33

1360.77

1347.49

kg

Mass after burn (Mf)

878.85

1053.93

1140.96

1129.82

kg

Gravity

9.81

9.81

9.805

9.805


deltaV (M/Sec)

630

630

630

630

M/sec

ISP

376

376

364.7

364.7

sec

Landing Prop/Take
-
off Capacity

0.278

0.258

0.236

0.248



Engine Parameters






Stage 1 Thrust/Weight Ratio

0.766

0.766

0.766

0.766


Stage 2 Thrust/Weight Ratio

0.268

0.268

0.268

0.268


Stage 1 Landing Thrust (kgf)

798.84

957.98

1042.60

1032.42

kgf

Stage 1 Launch Thrust (kgf)

669.59

865.29

1040.35

1012.10

kgf

Stage 2 Launch Thrust (kgf)

52.571

63.108

67.102

74.211

kgf

Engine Thrust/Weight Ratio

15.500

15.500

1
5.500

15.500


Stage 1 Engine Mass

43.199

55.825

59.100

57.400

kg

Stage 2 Engine Mass

3.392

4.071

3.410

3.770

kg


Vehicle Components






Capsule

20.0

20.0

20.0

20.0

kg

Cruise Stage

24.0

24.0

24.0

24.0

kg

Cruise Stage Fuel

4.0

4.0

4.0

4.0

kg

Sample M
ass

0.5

0.5

0.5

0.5

kg

Earth Return package (total)

48.5

48.5

48.5

48.5

kg

Project Rigel Mars Sample R
eturn


Kent Nebergall


17

MSR Project Rigel

Iter
.

1

Iter
.

2

Iter
.

3A

Iter
.

3B

Unit

Stage 2 Dry Stage

21.1

35.1

37.6

40.1

kg

Dry Total Stage 2 Mass

69.64

83.60

86.082

88.6

kg

Stage 2, Propellant + 1% waste

126.69

152.08

166.16

190.43

kg

Stage 2, Mass to Orbit

196.33

235.68

250.60

277.15

kg

Stage 1 Mass, Dry

87.97

131.72

174.00

166.70

kg

1 dry + 2 wet + payload

284.30

367.40

424.60

443.85

kg

Stage 1 Propellant + 1% waste

589.63

761.96

942.58

885.89

kg

Stage 1 Total Liftoff Mass

873.94

1129.36

1357.84

1320.97

kg

Stage 2 Wet + 1 Dry Mass

353.95

451.00

501.35

523.68

kg

Landing Stage
-

rover, H2

432.28

526.23

567.67

560.02

kg

Landing Gear, etc.

150.00

150.00

150.00

150.00

kg

Landing Stage Avionics

20.00

20.00

20.00

20.00

kg

ISRU Plant Mass (est. kg)

30.00

60
.00

60.00

60.00

kg

H2 Tank Thickness (m)

0.01

0.01

0.01

0.01

m

H2 Tank Material Density (kg/m3)

2700.00

2700.00

2700.00

2700.00

kg/m3

H2 Tank Exterior Volume

0.55

0.69

0.83

0.81

m3

H2 Tank Interior Volume

0.48

0.62

0.75

0.73

m3

H2 Dry Tank Mass (kg)

1
67.56

196.50

222.98

218.69

kg

Solar Array (est.)

64.73

99.73

114.69

111.34

kg

Rover

185.00

185.00

185.00

185.00

kg

Hydrogen Payload

34.31

43.78

53.11

51.56

kg

Landing Mass
-

landing propellant

878.85

1053.93

1140.96

1129.82

kg

Nominal Landing Propella
nt

163.78

196.40

219.81

217.66

kg

Nominal Total Landing Mass

1042.63

1250.33

1360.77

1347.49

kg

Max landing Propellant

589.63

761.96

933.24

877.12

kg

Max Landing Stage Allowed

2244.37

2121.92

2057.49

2010.89

kg

Nominal Entry Mass

1567.63

1775.33

1885.7
7

1872.49

kg

Aeroshell (backshell + Heat shield)

525.00

525.00

525.00

525.00

kg

MSL Entry Mass

2063.00

2063.00

2063.00

2063.00

kg

Cruise Stage

400.00

400.00

400.00

400.00

kg

Nominal Mass From Earth

1967.63

2175.33

2285.77

2272.49

kg

Maximum Mass from
Earth

6083.57

6214.58

6367.03

6260.81

kg

MSL Total Mass

2463.00

2463.00

2463.00

2463.00

kg

Ratio of Entry Mass

(MSR:MSL)

0.76

0.86

0.91

0.91


Ratio of Launch Mass

(MSR:MSL)

0.80

0.88

0.93

0.92


Margin from MSL Baseline

495.37

287.67

177.23

190.51

kg

F
uel Selected

Ethylene

Ethylene

Ethylene

Ethylene


Fuel Ratio

2.6 to 1

2.6 to 1

2.6 to 1

2.6 to 1


Total of Ratio

3.60

3.60

3.6

3.6


Fuel Ratio: Oxygen Part

2.60

2.60

2.6

2.6


Fuel Ratio: Propellant Part

1.00

1.00

1

1


Oxygen Fraction

0.722

0.722

0.722

0.722


Fuel Fraction

0.278

0.278

0.278

0.278


Oxygen Density Factor

1141.00

1141.00

1141.00

1141.00


Propellant Density Factor

567.92

567.92

567.92

567.92


Hydrogen Density Factor

70.97

70.97

70.97

70.97


LH2/Fuel Ratio

0.144

0.144

0.144

0.144


Rati
o in bold above shows the Iteration
3B

vehicle maintains an 8 percent margin below the
launch mass and a 9 percent margin below the entry mass of the Mars Science Lab.

Project Rigel Mars Sample R
eturn


Kent Nebergall


18

MSR Project Rigel

Iter
.

1

Iter
.

2

Iter
.

3A

Iter
.

3B

Unit


Stage 1 Propellant






Mass of Propellant (kg)

589.631

761.960

942.576

885.892

kg

M
ass of Oxygen (kg)

425.845

550.304

680.749

639.811

kg

Mass of Fuel (kg)

163.786

211.656

261.827

246.081

kg

Hydrogen Tank (m3)

0.483

0.617

0.748

0.726

m3

Stage 1 Oxygen (m3)

0.373

0.482

0.597

0.561

m3

Stage 1 Fuel (m3)

0.288

0.373

0.461

0.433

m3

Sphere

Dia Oxygen

0.893

0.973

1.044

1.023

m

Sphere Dia Fuel

0.820

0.893

0.958

0.939

m

Sphere Dia H2

0.974

1.056

1.126

1.115

m

Sphere Dia if LOX/Fuel in 1 sphere

0.974

1.056

1.126

1.115

m

Sphere Dia of Each Dual Tank

0.976

0.978

0.980

0.980

m


Stage 2 Propel
lant






Mass of Propellant (kg)

126.688

152.081

166.160

190.433

kg

Mass of Oxygen (kg)

91.497

109.836

120.005

137.535

kg

Mass of Fuel (kg)

35.191

42.245

46.156

52.898

kg

Stage 2 Oxygen (m3)

0.080

0.096

0.105

0.121

m3

Stage 2 Fuel (m3)

0.062

0.074

0.
081

0.093

m3

Combined LOX/Fuel (m3)

0.142

0.171

0.186

0.214

m3

Sphere Dia Oxygen

0.535

0.569

0.586

0.613

m

Sphere Dia Fuel

0.491

0.522

0.537

0.562

m

Sphere Dia
if

LOX/Fuel

in 1 sphere

0.648

0.688

0.709

0.742

m

Sphere Dia of Each Dual Tank

0.607

0.609

0.610

0.612

m


Landing Propellant






Mass of Propellant (kg)

163.779

196.404

219.810

217.664

kg

Mass of Oxygen (kg)

118.285

141.848

158.752

157.201

kg

Mass of Fuel (kg)

45.494

54.557

61.058

60.462

kg

Volume of Oxygen (m3)

0.104

0.124

0.139

0.138

m3

Volume of Fuel (m3)

0.080

0.096

0.108

0.106

m3


ISPP Demand






Total Propellant Needed

716.320

914.041

1108.737

1076.325

kg

Total Fuel Needed

198.978

253.900

307.982

298.979

kg

Total LH2 Needed to Produce

28.593

36.486

44.258

42.964

kg

LH2 Boil
-
off

Allowance

0.200

0.200

0.200

0.200


Total LH2 Mass Needed

34.312

43.783

53.109

51.556

kg

Total LH2 Volume Needed

0.483

0.617

0.748

0.726

m3

LH2 Tank Minor Axis

1.050

1.100

1.200

1.200

m3

LH2 Tank Major Axis

1.000

1.100

1.200

1.200

m3

LH2 Tank Z Axis

0
.900

1.000

1.000

1.000

m3

Actual Capacity of Tank

0.495

0.634

0.754

0.754

m3


Surface ISRU Production






Propellant Needed (kg)


716.32


914.04

1108.74

1076.32

kg

Surface Stay Time Allowed (Days)


500.00


500

500

500

days

Surface

Stay Time Allowed (Sols)



486.62

486.62

486.62

sols

Ave Prod Rate Per Day

1.433

1.828

2.217

2.153

kg

Project Rigel Mars Sample R
eturn


Kent Nebergall


19

MSR Project Rigel

Iter
.

1

Iter
.

2

Iter
.

3A

Iter
.

3B

Unit

Refined Prod Rate Per Day

1.407

2.110

2.562

2.487

kg

Refined Prod Rate Per Sol

1.446

2.168

2.493

2.420

kg

Power Demand of ISRU Unit (est)

809.
1

1
246.
6

1433.
6

1391.
7

watts/hr

Solar Array Mass (Zubrin)

64.727

99.728

114.690

111.336

kg

Solar Array Area (Zubrin) (Fixed)

21.576

33.243

38.230

37.112

m2

Power Demand Per 12 hr Sol

9709.
1

14959.
2

17203.
5

16700.
5

watts/sol

Peak MER Power Per Sol/M3

750

7
50

750

750

watts/m2

Current MER Power Per Sol/M3

310.67

310.67

310.67

310.67

watts/m2

Peak MER Power
-

Area Needed

12.945

19.946

22.938

22.267

watts/m2

Current MER Power
-

Area Needed

31.252

48.151

55.376

53.756

watts/m2

MER Equiv Power Average/Sol

530
.335

530.335

530.335

530.335

watts/m2

MER Ave Power Area Needed

18.307

28.207

32.439

31.490

m2

Actual Array Size
*

19.00

29.48

33.00

33.13

m2

Actual Array Peak Output
*


14,250


22,110


24,750

24,849

watts/sol

Actual Array Average Output
*



10,076


15,634

17,501

17,571

watts/sol

* This figure does not include the secondary array.


Solar
Array
Dimensions

Iteration 2

Iteration
3B

Lateral Array Panels (8)

1.1 X 3.2 m

1.1 X 3.64 m

First Inset Array


1.1 X 0.7 m

Second Inset Array


1
.1 X 0.3 m

Secondary Array

(Approximated from photos of Phoenix lander, from
which this array is derived.)

2 m dia.

2 m dia.



Iteration
3A

and
3B

Clarifications and Changes

Iteration
3A

and
3B

include more detailed data on Delta V requirements, and a ne
w pump
design developed at Laurence Livermore Labs specifically for MSR vehicles. A cursory
remediation for aerodynamic issues on ascent is also proposed. The difference between
Iteration
3A

and
3B

is that
3A

uses the ascent stage to circularize the orbi
t of the return
vehicle, whereas
3B

uses the Trans
-
Earth Injection stage to do this task. By doing the
staging slightly earlier, a small amount of mass was
reduced from the vehicle.


Iteration
3A

and
3B

also include more inefficiency allowances in the des
ign to move it
further from the ideal design and closer to something that would be realistically
constructed. It also does not allow for the velocity advantages of equatorial launch. This
allows future iterations to either
land at

the poles if the design

is found to be accurate or
land at

progressively lower latitude limits if the vehicle is found to be optimistic.



Delta V

The standard textbook Delta
-
V for Mars ascent to orbit is 4140 [6].

According to a more detailed study that specifically focuses o
n a 100
kg MAV that is liquid
-
fueled and accounts for the atmosphere (with
local speed of sound issues, etc.), that figure is actually 3900 for initial
takeoff and 257 for circularization, for a total of 4157. This is a fairly
Project Rigel Mars Sample R
eturn


Kent Nebergall


20

minor difference of 17 m/s.

However, the vehicle can then be
redesigned so that the Trans
-
Earth Injection (TEI) stage actually does
the circularization of the initial orbit, thus removing 240 m/s from the
Ascent stage and adding 257 to the TEI stage. The design
is

calculated
both w
ays to see if the increased demands on the first s
tage to carry the
larger second stage are offset by the lower final velocity required for
the first stage.


Propellant
Pumps

Work has been done at Laurence Livermore Labs in designing and
prototyping a 300

gram reciprocating fuel pump for MAV applications
[18]. This work assumes a lower specific impulse propellant (310s
versus 376 for Rigel), a lighter vehicle (100 kg versus roughly 1000 kg)
and far less engine thrust (102 kg versus 865 kg).
That said, th
e design
at this point calls for four pumps for the first stage and two for the
second, and each scaled up 40 percent to 500 grams. The dual
-
dual
pumps will allow two per side of the ascent stage. Using six identical
pumps rather than different designs f
or each stage should reduce
development costs.

It is assumed based on the original work that
2 percent of the propellant
will be burned to drive the pumps. After consulting with the author of
the paper, this was counted as a percentage against the ISP of t
he
engines.

This is partly offset by directing the exhaust in the direction
of the engine’s thrust. It may also be used to augment the RCS system.


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Project Rigel Mars Sample R
eturn


Kent Nebergall


21

Examining Configuration Issues

General Design Parameters

The go
al of this design will be maximum reliability at minimum cost with maximum
design margins where possible.
The configuration options will consume some of those
margins for the sake of reliability and cost.


Component

Description

Sample Return
Vehicle

This

is scaled to the minimum sample size


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W
here possible, MER, MSL, and other
then
-
flown
components are

reused with as little modification as possible.

2.

If a technology was developed for an alternate Mars vehicle
that was either canceled or not yet approved, that is
considered as the next option bec
ause some of the R&D cost
has already been spent.

3.

If something needs to be built from scratch, the physical
design will be kept as simple and foolproof as possible. The
general rule is that if it can’t be prototyped by a hobbyist in
瑨攠条牡来Ⱐ晩湤⁳潭整
桩湧⁴桡琠 a渮n




Project Rigel Mars Sample R
eturn


Kent Nebergall


22

The Vehicle Design

This section covers each design component in detail.

Sample
Return
Capsule

Overview

The capsule is a passive low density aeroshell which contains a core
sphere of crushable material. Within that is a three
-
chamber
c
anister
that contains
soil

and rock samples, a compressed atmospheric sample,
and an uncompressed atmospheric and filtered dust sample. The overall
mass is estimated at 20 kg plus
the
0.5 kg sample.


Background

Dr.
Zubrin’s original MSR discussion describ
ed a 0.5 kg sample
brought back in a 6 kg passive return capsule

[4]
.
D
etailed papers for
the ESA from EADS discuss a similar design, but scale it up to a 5 kg
sample capsule contained in a 60
-
80 kg aeroshell

[8]
.


Comparison of
EADS and this
MSR Desig
n

Component

EADS Design

Rigel Design

Sample
Canister

5 kg with sample

1 kg with sample

Sample
Canister
Diameter

25 cm exterior

9 cm interior,

10.3 cm exterior

Overall Mass with Sample

60
-
80 kg

20 kg

Aeroshell Diameter

140 cm

50 cm

Crumple zone di
ameter

43 cm

30 cm

Cone Angle

45 degrees

45 degrees

Sample Container Shape

Sphere

Cylinder


Diagram

11.0cm.
50.0cm.
Aeroshell
Crumple Zone
Sample
30.0cm.


Differences

between Rigel
and EADS



The sample capsule is a cylinder rather than a sphere to simplify
loading and to allow atmospheric and dust sa
mple sections.




To give control of weight and balance, and also to amplify the
protection of the crumple zone, the vehicle has a disc that fits
behind the sample capsule toward the center of the crumple zone.
This is also the mount point for fixing the sa
mple capsule in place.


Project Rigel Mars Sample R
eturn


Kent Nebergall


23

General Tech
Level

P
assive, low
-
density capsules arriving at high entry velocities are
unprecedented
. T
his is the technology of choice for the Russian
Phobos
-
Grunt

mission currently being planned

[15]
.


The vehicle itself

(w
ith
its lack of moving parts, very simple
electronics, and well
-
understood aerodynamics
)

require
s

minimal
testing. Early sample returns from the Lunar south pole or other
missions that may follow Stardust or Genesis in deep space sample
return should standard
ize on a given passive aerodynamic capsule
design

where appropriate
.


Soil and Rock
Sample
Container


The
soil
sample
section of this

the
canister

is a
9

cm interior diameter,
6

cm long titanium
chamber

containing a stack of soil collection tapes
from the

rover.
Th
e rover is equipped with a reel
-
to
-
reel

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Project Rigel Mars Sample R
eturn


Kent Nebergall


24

Since the
soil and rock chamber of the canister
is em
pty
during most of
the surface

time, the input valve for
the filter

is at the very back of the
interior cylinder. This

reduces the number of openings on the outside
of the later
-
sealed canister and also
prevents the
filtered
dust from
pooling on one side
and potentially leaving the capsule less balanced.


Samples

from this section will be critical to the design of future air
filters for crewed vehicles and other robotic ISPP systems.


Sealing
Mechanisms

To seal the soil sample cylinder, the cylinder c
ontains two
Teflon tape
coated threaded elements that are screwed together. Since the threads
are filled with tape before launch, they cannot be clogged with dirt or
dust prior to the threading operation. The tape itself on both sides
would be covered wi
th a thin plastic sheet that could be pulled away
with a drawstring just prior to sealing.


Valves for the air samples are closed either mechanically or
pyrotechnically in the same manner as a cable cutting g
uillotine

for
separating sections of a spacecraf
t. After closing, a leak test would be
conducted and a secondary closing mechanism fired or closed if
necessary. Finally, a third pin mechanism would break a glass vial
containing an expanding gap filler between the two valves to ensure
that the air samp
les are properly sealed even if an impact or other even
t

jars the valves.


Crumple zone

Surrounding th
e

sample cylinder is the “crumple zone” of low density
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Project Rigel Mars Sample R
eturn


Kent Nebergall


25

Entry TPS
Shield

The outer entry cone is a 45 degre
e, 3 cm thick zone. The overall
diameter of the entry capsule is 50 cm.
It is made of an ablative
material with a high density surface and a low density interior.


Structure

A series of composite
spars and ribs bond the entry shield to the
crumple zon
e

core. A second skin seals the back of the assembly and
contains the electronics in small, thinly
-
insulated composite boxes.
The beacon and strobe are on opposite sides for balance, and two
passive transponders are offset from both 90 degrees. Gaps wit
hin the
structure are filled with very low density foam for added shock
insulation on impact.


Mass

The estimated mass, including the sample, was originally calculated at
6.5 kg.
T
o be conservative it will be listed as a total mass of 10 kg
including a
ny structural connections.


Electronics

This consists of a passive transponder, an active radio beacon, and a
strobe light

on a timer
.

The radio beacons are activated upon departure
from the cruise stage. The strobe is activated
shortly before

impact.
The batteries are charged from the cruise stage just before
disconnection.


A pressure sensor exists in the atmospheric and the soil sample
sections. This alerts the cruise
stage

if a leak occurs. This may result
in rotating the vehicle relative to the s
olar UV light to allay fears of
back contamination.
In an extreme case, the leak may result in
diverting the capsule away from
E
arth and avoiding the issue.


Operations:

Loading the
Capsule

The entire sample capsule cylinder is kept in the sample loading

bay
prior to
launch
. The loading sequence is as follows:


1.

During the Martian Summer, the air dust sample section is sealed.

2.

The air sample section, after pressurization, is
also sealed during
the Martian Winter. This gives a sample of the atmosphere in

opposite seasons for comparison.

3.

The capsule is unsealed from its sterilization bag by use of a draw
-
string. This is pulled
away

from two directions (either will work)
in a
manner

similar to the arm on the Phoenix lander
.

4.

The rock sample bag is loaded at

the bottom of the sample cylinder,
followed by the tape samples

and the core sample
.

5.

The cylinder
is

sealed

with a threaded cap
.

6.

The cylinder is checked for center of gravity in all three axes.

7.

The loading manipulator both adjusts the balance disk in th
e
Crumple Zone to offset the CG within the cylinder, and sets the
screw
-
in fitting to be in the exact center of the loading bay for
mating with the sample cylinder.

Project Rigel Mars Sample R
eturn


Kent Nebergall


26

8.

The sample cylinder is moved into the capsule bay by the
manipulator and screwed the approp
riate number of turns into the
move the CG forward to the appropriate location.

9.

The cylinder is capped with an additional crumple z
one/insulation
section and the outer skin door is closed and locked
.
The insulation
is also locked into place within the low

density crumple zone with a
soft shallow dovetail joint.

10.

The Cruise Stage, which has been elevated 1
5

cm

after landing, can
be

lowered into place if appropriate.


Operations:

Capsule Entry

The capsule entry sequence is as follows
:


1.

The cruise stage charg
es the capsule battery and runs final tests of
the capsule electronics. The capsule may be close enough to
E
arth
that the beacon may be detected from ground tracking.

2.

The
c
ruise stage ensures the capsule is correctly oriented and
transmits final coordinat
es to the ground.

3.

The cruise stage calculates the appropriate delay and sets a timer on
the capsule to sequence the activation of the strobes on the capsule.

4.

The electronic coupler between the capsule and cruise stage is
severed.

5.

The capsule is spun
fast
er
to ensure stability using the cruise stage
RCS system.

6.

A spring
-
loaded mechanism on the cruise stage gives final
separation.

7.

The capsule maintains the beacon but not the strobes until
atmospheric entry.

8.

When the timer completes, several km above the gro
und, the
strobes become active. This simplifies tracking from aircraft and
ground crews.

9.

The capsule makes landfall.

10.

A chase crew in a helicopter

switches off the electronics,

places the
capsule in a secure container and loads it for return to the
laborat
ory.


Planetary
Protection
(Earth to Mars)

Before leaving Earth, t
he entire capsule is sterilized by
autoclave
. The
entire capsule section is sealed in plastic

that was part of the autoclave
process.


Planetary
Protection
(Mars to Earth)

Since the TPS s
ection has a high probability of damage on impact, the
core of the structure is filled with foam that is biologically shielded by
the overall skin from and Martian contamination. Since pieces of this
foam may be scattered on impact, it is best they have n
ever touched the
Martian atmosphere or dust. Filling this space will a low density foam
not only gives those contaminants no where to go but adds to the
shock

absorption on impact.

Project Rigel Mars Sample R
eturn


Kent Nebergall


27

Optional
Equipment
:
Separation and
Entry Camera

It may be possible to
inc
lude

a rear
-
facing rocket camera that would be
activated on separation, record video of the cruise stage departure
,
entry, and landing

to solid state memory before recovery. This video
would be an added public relations bonus but also provide visual data
on the stability, plasma sheath formation, and other dynamics of the
entry

for use in future designs
.




Project Rigel Mars Sample R
eturn


Kent Nebergall


28

Earth Return Cruise Stage


Overview

The cruise stage is a fairly small RCS, navigation, and communication
platform that handles the return capsule fr
om TEI until just before
Earth atmospheric entry. The general arrangement has a pair of
hydrazine tanks at opposite sides of the stage with a four
-
engine RCS
pod at the end of each tank. At 90 degrees from the tanks are two
folding rectangular solar pane
ls. The center of the vehicle contains the
electronics, small high
-
gain antenna, and the mount for the capsule. A
low gain antenna is also placed along each solar panel for
communications when close to Earth and when the vehicle is facing the
opposite di
rection. A small radiator for the electronics is contained
between the capsule and the body to ensure it is in the shade in most
orientations.


Layout:

Side View with
capsule access
exposed

RCS Pod (2)
Antenna
Sample Capsule
Electronics, Solar Panel (f olded)

Layout:

Front View

(Solar Panels
Deployed)

Solar Panel (2)
Antenna
Sample Capsule

Illustration:

W
ings
deployed,
wings folded,
and sample
canister


Project Rigel Mars Sample R
eturn


Kent Nebergall


29

Inverted
Configuration

The entire entry vehicle and cruise stage configuration is flipped upside
down for the following reasons…



It e
nables the antenna and solar panels to deploy at the surface or in
Mars

orbit.



It c
ompletely protects the entry shield from exposure while on the
surface.



It r
educes the overall height of the vehicle by 22 cm.


Dimensions


Mass

24 kg plus 4 kg hydrazine propellant


Dimensions (folded, with
capsule)

8
0

cm (across RCS Pods)

30 cm (across capsule)

10
2

cm (with capsule raised for
sample access)

8
7

cm (with capsule lowered into
flight configuration)


Dimensions (deployed)

300 cm (across panels)

8
0

cm (across RCS Pods)


Electronics Box

30 cm cube


RCS Pod: Tank Diameter

18
cm sphere (2)


RCS Pod: Engine Cluster

10 cm across, 5 cm thick (2),

4 engines per cluster


High Gain Antenna

28 cm (same as MER

[7]
)


Solar Array Dimensions

30 cm wide, 120 cm long each


Solar Array Output

154 watts at Earth

77 watts at Mars


Ba
ttery

10 amp hour Li
-
Ion (MER)



General Tech
Level

The smallest cruise stage launched has been the Mars Polar Lander
cruise ring, which at 56 kg (82 kg with the Deep Space 2 capsules and
equipment) was able to power and guide a 494 kg capsule

[5]
.


T
echnology for this return vehicle is fairly similar with two exceptions.
First, Mars missions tend to use the payload’s computers, whereas the
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Project Rigel Mars Sample R
eturn


Kent Nebergall


30

also be used to make minor changes to the center of gravity before
takeoff for proper launch. It consists of four columns set on motorized
floating screws, with a flexible sheath around the assembly to keep dust
out during the surface stay. If one screw jams, the other three can
break a failsafe and force the other shaf
t to move. If this mechanism is
judged too complex, it could simply be moved pneumatically during
the loading operation and locked down mechanically before takeoff.
The columns are supported laterally by a

mounting

ring, and this ring
also connects the cr
uise stage to the TEI stage.
The cruise stage is the
section raised and lowered, with the mounting ring and lander kept in
place.


Navigation

A star and sun tracker arrangement is used, much as they are for
journeys to Mars. Navigation information based

on ground tracking
can also be uplinked to the computer on board, especially as the vehicle
approaches Earth.


Note that the computer onboard the cruise stage is also in charge of
guidance of the entire vehicle stack during
o
utbound
c
ruise, Mars
descent,
Mars ascent, and Trans Earth Injection.

The lander computer
acts as a backup during the outbound trip.


RCS System

The RCS system for the cruise stage allows navigation en route to
Earth, placement of the capsule in an Earth entry path, and placement
of
the cruise stage in either an Earth entry path for destruction or an
Earth flyby path for disposal at escape velocity.


The general arrangement is a core of two spherical titanium tanks
containing hydrazine and placed in opposite locations on the outside

of
the platform. This allows the spin of the vehicle (2 RPM) to provide
propellant to twin RCS pods


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Project Rigel Mars Sample R
eturn


Kent Nebergall


31

flight with the lowest solar input

with a medium gain antenna
, whereas
the communications demands for this vehicle

when it reaches Earth

are
over much shorter distances and with far more
solar
power per square
meter
.


Perpendicular to the line containing the RCS tanks are the two
rectangular solar panels. Both side
s of both panels have solar cells, and
they will be able to provide minimal power even if folded
(
provided
they are not shaded
)
. These deploy from opposite sides and reach
beyond the shadow of the return capsule.

They can even be
temporarily deployed on
the surface if necessary.


It the vehicle goes into safe mode, the panels are canted at a 20 degree
angle relative to each other. If both panels are in a single line, and the
sun is edge
-
on, neither panel will receive power and the probe will
eventually b
e lost. By canting them back at an angle, this breaks up
this single line and at least one side of one panel will be illuminated
regardless of spacecraft orientation. The battery can then be slowly
charged until it has enough power to communicat
e

with Ea
rth.


The solar panels are folded into four sections and measure 120 by 30
cm each

when deployed
. This results in 154 watts near Earth and 77
watts near Mars when the panels are properly aligned. To simplify
power handling, the battery is disproportionat
ely large and the system
as a whole
can

run in several power modes depending on current
demand. The battery is slowly charged between communications and
other demanding bursts to
compensate for

the small panels.


Telemetry

The main
antenna

is 28 cm in
diameter and is a direct copy of the MER
antenna

[7]
. It is the main antenna for both the return cruise and
surface operations.


During the return voyage, since the antenna sits on the axis of rotation,
tracking an offset target (
E
arth) while rotating sim
ply involves rotating
the dish on one degree of freedom.
The antenna is probably equipped
with a radome to minimize transonic stress on ascent and to minimize
dust contamination on the surface.


Planetary
Protection
(Mars to Earth)

After separation fro
m the capsule,
and assuming there is insufficient
fuel to miss Earth,
the cruise stage uses its remaining fuel to place itself
in a tumble to ensure the entire structure is directly exposed to and
consumed by
atmospheric

entry plasma. The lightweight stru
cture
should be consumed completely, and the titanium tanks outer surface
insulation will burn away

even if the tanks are empty and do not
explode
. The cruise stage interior is sealed against dust contamination
and therefore will not harbor any potential
Martian organisms.

Project Rigel Mars Sample R
eturn


Kent Nebergall


32

T
rans
-
Earth Injection (T
EI
)

Stage

(Second Stage)


Overview

Since the stages are nested, the RCS
p
ods are close to the center of
gravity for the two stage configuration and
steer

from the rear like fins